Real-time aircraft status detection system and method

ABSTRACT

A low power method for determining whether a cargo destined for air transport is in a flying state having the steps of: providing a housing for attachment to a cargo the housing having: an accelerometer for detecting a linear acceleration, a gyroscope for detecting an angular rate, a controller measuring a linear acceleration with the accelerometer, measuring an angular rate with the gyroscope, providing the measured linear acceleration and angular rate to the controller, and generating a flight status output signal indicating whether the housing is in a flying state as a function of the linear acceleration signal and angular rate signal.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority to U.S. provisional patent applicationNo. 61/606,364 which was filed on Mar. 2, 2012, which is herebyincorporated by reference in its entirety.

TECHNICAL FIELD

The technical field relates to aircraft status detection, includingaircraft takeoff, landing, flying, and aircraft proximity detection.

BACKGROUND ART

Aircraft flight status detectors are generally known. For example, U.S.Pat. No. 7,791,455 entitled “Method and Apparatus for AutonomousDetection of a Given Location or Situation” is directed to system fordetecting the presence of operational aircraft through detecting thepresence of a predetermined signal, the arrival at a given location, anda change in air pressure.

BRIEF SUMMARY OF THE INVENTION

With parenthetical reference to the corresponding parts, portions orsurfaces of the disclosed embodiment, merely for the purpose ofillustration and not by way of limitation, provided is a low powermethod for determining whether a cargo (114) destined for air transportis in a flying state having the steps of: providing a housing (111) forattachment to a cargo the housing having: an accelerometer (117) fordetecting a linear acceleration; a gyroscope (118) for detecting anangular rate; a controller (110); measuring a linear acceleration withthe accelerometer; measuring an angular rate with the gyroscope;providing the measured linear acceleration and angular rate to thecontroller; and generating a flight status output signal (115)indicating whether the housing is in a flying state as a function of thelinear acceleration signal and angular rate signal.

The method may further have the step of shutting off a radio as afunction of the flight status output signal. The measured linearacceleration hay have a three independent axis linear accelerationsignal. The measured angular rate may have a three independent axisrotation rate signal. The measured angular rate signal may have asampling rate of about 25 Hz.

The flight status output signal may have an in-flight status state andan on-the-ground status state. The in-flight status state may have averify-takeoff status state. The method may further have the step ofproviding a radio turn off signal when the flight status output signalmay have an in-flight status. The function may be a state machinealgorithm.

The method may further have the step of powering the gyroscope as afunction of the flight status output signal. The gyroscope may bepowered off when the flight status output signal may have anon-the-ground status state. The gyroscope may be powered on when theflight status output signal may have a verify-takeoff status state. Thefunction may have the steps of: filtering the measured linearacceleration to produce a filtered linear acceleration; filtering themeasured angular rate to produce a filtered angular rate; and changing astate as a function of the filtered linear acceleration and filteredangular rate.

The step of generating a flight status output signal may have the stepof removing a gravitational component of the measured linearacceleration. The step of removing a gravitational component may have ahigh pass filter. The high pass filter may have a cutoff frequency ofabout 0.01 Hz. The method may further have the step of saving theremoved gravitational component.

The step of generating a flight status output signal further may havethe step of detecting a linear acceleration characteristic of anaircraft takeoff. The linear acceleration characteristic of takeoff isan acceleration signal with a frequency between 0.01 Hz and 0.1 Hz and amagnitude of about 0.2 g. The linear acceleration characteristic of atakeoff further may have an acceleration signal which may have amagnitude between about 0.15 g and 0.5 g maintained for a duration ofabout two seconds. The step of detecting a linear accelerationcharacteristic of an aircraft takeoff further may have a step of lowpass filtering.

The method may further have the step of saving a takeoff acceleration ofthe linear acceleration signal when the linear accelerationcharacteristic of an aircraft takeoff is detected. The saved takeoffacceleration may have a windowed average. The step of generating aflight status output signal further may have the step of changing froman on-the-ground state to a verify-takeoff state when the linearacceleration characteristic of an aircraft takeoff is detected. The stepof saving a gravitational acceleration component from the linearacceleration signal.

The gravitational acceleration is saved from a time period before atransition to a verify-takeoff state. The time period is about 7seconds. The step of generating a flight status output signal furthermay have the step of detecting an angular rate characteristic of anaircraft takeoff. The angular rate characteristic of an aircraft takeoffis an angular rate in which an x-axis angular rate, a y-axis angularrate, and a z-axis angular rate are all less than an angular rate limitthreshold.

The measured angular rate is transformed through a rotation, therotation configured and arranged to reorient the measured angular rateinto a yaw rate, a roll rate, and a pitch rate. The rotation is a crossproduct of the saved gravitational acceleration and the saved takeoffacceleration. The method may further have the step of integrating theangular rate with respect to time to produce a pitch angle displacement,a yaw angle displacement, and a roll angle displacement.

The method may further have the step of setting the flight status outputsignal to an on-the-ground state when the pitch angle displacement, yawangle displacement, or roll angle displacement are greater than an angledisplacement threshold.

The method may further have the steps of: providing a pressure sensor.The method may further have the step of detecting when the pressuresensor may have a pressure decrease rate less than a takeoff pressuredecrease threshold. The method may further have the step of detectingwhen the pressure sensor may have a pressure increase rate greater thana landing pressure increase threshold.

In another aspect, provided is a system for detecting an aircraft flightstatus having: a housing having: an accelerometer for detecting a linearacceleration and having a linear acceleration output signal; a gyroscopefor detecting an angular rate and having an angular rate output signal;and a controller configured and arranged to produce a flight statusoutput signal as a function of the linear acceleration output signal andthe angular rate output signal, the flight status output signal havingan in-flight state and another state.

The controller is further configured to shut off a radio as a functionof the flight status output signal. The linear acceleration outputsignal may have a three independent axis linear acceleration signal. Theangular rate output signal may have a three independent axis rotationrate signal. The angular rate output signal may have a sampling rate ofabout 25 Hz. The flight status output signal may have an on-the-groundstatus state. The in-flight status state may have a verify-takeoffstatus state.

The controller may be configured and arranged turn off the radio whenthe flight status output signal may have an in-flight status. Thefunction is a state machine algorithm. The controller is configured andarranged to power the gyroscope as a function of the flight statusoutput signal. The controller is configured and arranged to power offthe gyroscope when the flight status output signal may have anon-the-ground status state.

The controller is configured and arranged to power on the gyroscope whenthe flight status output signal may have a verify-takeoff status state.The function may have a filter configured and arranged to produce afiltered linear acceleration, and a filter configured and arranged toproduce a filtered angular rate. The controller is configured andarranged to remove a gravitational component of the linear accelerationoutput signal. The system may further have a high pass filter configuredand arranged to remove the gravitational component from the linearacceleration output signal. The high pass filter may have a cutofffrequency of about 0.01 Hz.

The controller may be configured and arranged to save the removedgravitational component. The controller may be configured and arrangedto detect a linear acceleration characteristic of an aircraft takeoff.The linear acceleration characteristic of takeoff is an accelerationsignal with a frequency between 0.01 Hz and 0.1 Hz and a magnitude ofabout 0.2 g. The linear acceleration characteristic of a takeoff furthermay have an acceleration signal which may have a magnitude between about0.15 g and 0.5 g maintained for a duration of about two seconds.

The system may further have a low pass filter configured and arranged tofilter the linear acceleration output for detection of an accelerationcharacteristic of an aircraft takeoff. The controller may be configuredand arranged to save the acceleration output signal when a linearacceleration characteristic of an aircraft is detected. The savedacceleration output signal may have a windowed average.

The controller may be configured and arranged to change the flightstatus output signal from an on-the-ground state to a verify-takeoffstate when the linear acceleration characteristic of an aircraft takeoffmay be detected. The controller may be configured and arranged to save agravitational acceleration component from the linear accelerationsignal.

The gravitational acceleration may be saved from a time period before atransition to a verify-takeoff state. The time period may be about 7seconds. The controller may be configured and arranged to detect anangular rate characteristic of an aircraft takeoff.

The angular rate characteristic of an aircraft takeoff may be an angularrate in which an x-axis angular rate, a y-axis angular rate, and az-axis angular rate are all less than an angular rate limit threshold.The system may further have a rotation transformer configured andarranged to transformed the angular rate output signal through arotation, the rotation configured and arranged to reorient the measuredangular rate into a yaw rate, a roll rate, and a pitch rate. Therotation may be a cross product of the saved gravitational accelerationand the saved takeoff acceleration.

The system may further have an integrator configured and arranged tointegrate the angular rate with respect to time to produce a pitch angledisplacement, a yaw angle displacement, and a roll angle displacement.The controller may be configured and arranged to set the flight statusoutput signal to an on-the-ground state when the pitch angledisplacement, yaw angle displacement, or roll angle displacement aregreater than an angle displacement threshold.

The system may further have a pressure sensor. The controller may beconfigured and arranged to detect when the pressure sensor may have apressure decrease rate less than a takeoff pressure decrease threshold.The controller may be configured and arranged to detect when thepressure sensor may have a pressure increase rate greater than a landingpressure increase threshold.

In another aspect, provided is a method of determining an aircraftflight status state among a set of flight status states having the stepsof: providing a housing for attachment to a package the housing having:an accelerometer for detecting a linear acceleration and having anacceleration output signal; a gyroscope for detecting a rotationalacceleration and having an angular acceleration output signal; receivingan acceleration output signal from the accelerometer; receiving anangular rate output signal from the gyroscope; determining the aircraftflight status state as a function of the acceleration output signal andthe angular rate output signal,the set of flight status states having anin-flight state and another state.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is an object diagram of a first embodiment aircraft flight statusdetection system incorporated into a cargo tracking device.

FIG. 2 is an algorithm and state diagram of the flight status detectionsystem shown in FIG. 1.

FIG. 3. is a flow diagram of the on-the-ground state processing in thesystem shown in FIG. 1.

FIG. 4. is a flow diagram of the verify-takeoff state processing in thesystem shown in FIG. 1.

FIG. 5 is a state diagram of a pressure sensing state machine foranother embodiment flight status detection system.

DESCRIPTION OF THE PREFERRED EMBODIMENTS

At the outset, it should be clearly understood that like referencenumerals are intended to identify the same structural elements, portionsor surfaces consistently throughout the several drawing figures, as suchelements, portions or surfaces may be further described or explained bythe entire written specification, of which this detailed description isan integral part. Unless otherwise indicated, the drawings are intendedto be read (e.g., cross-hatching, arrangement of parts, proportion,degree, etc.) together with the specification, and are to be considereda portion of the entire written description of this invention. As usedin the following description, the terms “horizontal”, “vertical”,“left”, “right”, “up” and “down”, as well as adjectival and adverbialderivatives thereof (e.g., “horizontally”, “rightwardly”, “upwardly”,etc.), simply refer to the orientation of the illustrated structure asthe particular drawing figure faces the reader. Similarly, the terms“inwardly” and “outwardly” generally refer to the orientation of asurface relative to its axis of elongation, or axis of rotation, asappropriate.

The disclosed flight status detection system is a system for detectingaircraft events, such as takeoff and landing events, through the use ofaccelerometers, gyroscopes, and/or other sensors. The flight statusdetection system can be used together with a device having a cellularradio in order to turn off the cellular radio during a takeoff event andturn back on upon landing.

For example, the flight status detection system can be used within aflight cargo tracking device having a cellular data modem. While thecargo tracking device is on the ground, periodic tracking reports aresent over the cellular modem to a remote server. When the flight statusdetection system detects that the tracking device is on an aircraft thatis taking off, the cellular modem of the cargo tracking system is turnedoff. This will prevent the cellular modem radio from interfering withaircraft radios and sensors. Such a system is useful in order to complywith flight safety regulations that require that radios be shut offprior to takeoff. When the flight status detection system detects thatthe aircraft has landed, the cellular modem is turned back on, allowingthe cargo tracking device to resume sending out periodic trackingreports.

Referring now to the figures, and more particularly to FIG. 1, disclosedis flight status detection system 110 configured and arranged withincargo tracking and reporting module 111. Tracking module 111 istypically affixed to or placed within a cargo container, such as cargocontainer 114. Tracking module 111 gathers and transmits transportationdata 132 to remote computer 120 while cargo container 114 istransported. The major components of tracking module 111 are flightstatus detection system 110, accelerometer 117, gyroscope 118, and radio113.

Accelerometer 117 is a low power three axis MEMS linear accelerometersuch as an LIS3DH from STMicroelectronics of Geneva Switzerland.Accelerometer 117 provides three axis linear acceleration data A (Ax,Ay, Az) 120 to flight status detection system 110. Gyroscope 118 is alow power MEMS gyroscope such as a CMR3000, from VTI Technologies (nowMurata Manufacturing Oy of Nagaokakyo, Kyoto). Gyroscope 118 is turnedon and off through on/off command 122 from system 110. Gyroscope 118provides three axis rotation rate data R (d⊖x, d⊖y, d⊖z) 121 to system110.

System 110 is implemented on a microcontroller having its own flashmemory, such as an Atmel AVR Atmega328. System 110 contains softwarealgorithm and state machine 125. System 110 provides on/off command 115to radio 113. System 110 is also connected to radio 113 through datatransfer line 133. with Radio 113 is a radio such as a GE 865-QUADGSM/GPRS from Telit Wireless Solutions, Trieste, Italy. Radio 113 iscapable of transmitting and receiving data to/from a remote locationsuch as a cellular tower connected to the Internet over wireless link132.

FIG. 2 is a state machine diagram of algorithm 125 running in flightstatus detection system 110. Algorithm 125 generally receives inputs ofaccelerometer data (A) 120 and gyroscope data (R) 121; and has outputsGYRO_ON/OFF_OUTPUT 122 and RADIO_ON/OFF_OUTPUT 115. Algorithm 125contains a state machine which generally has three separate statesincluding: GROUND_STATE 151, VERIFY_STATE 153, and FLYING_STATE 155. Aseries of variables 160 and parameter constants 170 are used byalgorithm 125.

Accelerometer data 120 is a stream of time sampled acceleration vectorvalues having an acceleration sample for each dimension (Ax, Ay, Az).Acceleration data is sampled at a rate of 20 Hz. Based upon power andaccuracy tradeoffs, acceleration data may be sampled at much higherrates, such as up to 10 kHz. Note that there is no guarantee thataccelerometer 117 has any of its axes (x, y, z) aligned with thedirection of gravity when cargo 114 is stowed on an aircraft, truck, orwarehouse. In other words, cargo 114 may be a box oriented on an anglesuch that neither of Ax, Ay, nor Az give a measurement along a worldz-axis (an axis aligned with the direction of gravity). Gyroscope data121 is also a series of three dimensional vector values (Rx, Ry, Rz).Each of Rx, Ry, and Rz represent the rotation rate along threeindependent axes. Gyroscope data is also sampled at a rate of 20 Hz,however, higher sampling rates may be used when power is not a highconcern. Alternative gyroscopes may be used which provide angularacceleration, jerk, velocity, and/or position, or any combinationthereof.

The first algorithm state, GROUND_STATE 151 is both a starting state aswell as a state corresponding to a flight status of when system 110believes the system is not on an aircraft taking off or flying. InGROUND_STATE 151, GYRO_ON/OFF_OUTPUT 122 is set to OFF (i.e. lowvoltage, or FALSE) causing gyroscope 118 to be off. Gyroscope 118 willthus not require operating power in this state. Also in GROUND_STATE151, RADIO_ON/OFF_OUTPUT 115 is set to ON (i.e. low voltage HIGH, orTRUE). Therefore, in this state, radio 113 is able to transmit data 132.

In GROUND_STATE 151, accelerometer data 120 is read and analyzed. If anaccelerometer data sequence that is characteristic of a potentialaircraft takeoff is identified, recent accelerometer data is saved, andthe algorithm state machine enters VERIFY_STATE 153.

Once in VERIFY_STATE 153, GYRO_ON/OFF_OUTPUT 122 is transitioned fromOFF to ON (low to high voltage, FALSE to TRUE). This causes gyroscope118 to turn on and start generating gyroscope data R 121. InVERIFY_STATE 153, current gyroscope data R 121 and accelerometer data A,120 is analyzed and compared to the accelerometer data saved at thepotential takeoff. Several tests are done with the gyroscope data andaccelerometer data over time. If the analysis indicates that a takeoffevent has not occurred, the state is returned to GROUND_STATE. However,if the analysis indicates that a takeoff has indeed occurred, the statemachine enters FLYING_STATE 155.

In FLYING_STATE 155 both RADIO_ON/OFF_OUTPUT 115 and GYRO_ON/OFF_OUTPUT122 are set to OFF (voltage low, FALSE). Thus, in this state, neithergyroscope 118 nor radio 113 consume power, and radio signals from radio113 which could potentially interfere with flight instruments areprevented. Accelerometer data 120 is still sampled and is analyzed for apattern that is characteristic of an aircraft landing. If anaccelerometer pattern characteristic of an aircraft landing is sensed,the state machine reenters GROUND_STATE 151.

Upon reentering GROUND_STATE 151 the algorithm sets RADIO_ON/OFF_OUTPUT115 to ON (voltage high, TRUE). The algorithm will then continue toprocess data and step through the state machine as described above.

FIG. 3 is a datapath diagram of the processing done in GROUND_STATE 151.In GROUND_STATE 151, RADIO_ON/OFF_OUTPUT 115 is ON allowing radio 113 tofunction, and GYRO_ON/OFF_OUTPUT 122 is OFF, preventing gyroscope 118from using power. The algorithmic processing in GROUND_STATE uses thestream of accelerometer data 120 to make the decision of whether acandidate takeoff signature has been sensed. Accelerometer data 120 isfirst filtered by gravity filter 181 to produce gravity filter output182. Gravity filter 131 is a first order low pass Butterworth filterwith a cutoff frequency of 0.01 Hz. Since cargo 114 is typically notrotated slowly over a one hundred second interval, the filteredaccelerometer value 182 coming out of gravity filter 131 is typically agood representation of the linear acceleration experienced byaccelerometer 117 due to gravity. In other words, the resulting vector182 is likely to have an orientation pointing straight up relative tothe Earth. Vector 182 is passed through normalization block 183 whichscales the magnitude of vector 182 to unity length to produce gravityvector G 184. Gravity vector G 184 is an important variable which isused in multiple functions of algorithm 125.

Referring back to gravity filter output 182, output 182 is subtracted185 from original accelerometer sample 120 to produce vector A_horiz186. A_horiz 186 is a vector which is equivalent to the originalacceleration A 120 but with the gravity vector component removed. Inother words, A_horiz represents the horizontal, or perpendicular togravity (tangent to Earth) component of acceleration A 120. For example,if cargo 114 is accelerated along a runway, A_horiz 186 will generallyrepresent the component of acceleration in the direction along therunway.

A_horiz 186 is then passed through runway filter 187 to produce output188. Runway filter 187 is a first order Butterworth low pass filter witha frequency cutoff of 0.5 Hz. Runway filter 187 generally removes highfrequency accelerations which are characteristic of bumps on a road,rapid automotive accelerations due to speeding up, braking, or turning,and/or cargo handling shocks to yield an acceleration vector having onlythe targeted frequencies. Output 188 is then passed through magnitudeblock 189 to produce A_target_horiz_magnitude 190, which is merely thevector length (magnitude) of output 188.

A_target_horiz_magnitude 190 represents the magnitude of the originalacceleration after the gravity component is removed and the unwantedhigh frequency components are removed. In other words,A_target_horiz_magnitude 190 represents the magnitude of the horizontalacceleration in the targeted frequency range. The calculations ofA_target_horiz_magnitude and G are accurate regardless of theorientation that cargo 114 is in. If A_target_horiz_magnitude is a valuethat is typical of an aircraft takeoff, the state is changed toVERIFY_STATE 153. More specifically, if A_target_horiz_magnitude isgreater than threshold A_THRESHOLD_MIN 171 and less than thresholdA_THRESHOLD_MAX 172, then the state machine will enter VERIFY_STATE 153.Before transitioning to VERIFY_STATE 153, some variables are stored forlater use. Gravity Vector G 184, is stored to G_takeoff 161, and A_horiz186 is stored in A_horiz_takeoff 162. Instead of saving the most recentvalues of G 184 and A_horiz 186, it is beneficial to store the values ofG 184 and A_horiz 186 from 2 to 8 seconds prior to the detected takeoffevent.

In VERIFY_STATE 153, gyroscope 118 is turned on by settingGYRO_ON/OFF_OUTPUT 122 to ON. Generally, RADIO_ON/OFF_OUTPUT 115 is notchanged and remains ON, however, it may be beneficial in someimplementations to turn RADIO_ON/OFF_OUTPUT OFF 115 in the VERIFY_STATE153. The A_target_magnitude 190 is calculated in VERIFY_STATE 153 justas it was in GROUND_STATE 151. If at any time in VERIFY_STATE 153A_target_magnitude 190 drops below A_THRESHOLD_MIN 171 or rises aboveA_THRESHOLD_MAX 172, the state is returned to GROUND_STATE 151. In otherwords, if the magnitude of the horizontal acceleration in the targetedfrequency range is either too high or too low to be characteristic of anaircraft takeoff, the algorithm determines that a takeoff event has notoccurred and returns the state back to GROUND_STATE 151.

The processing of accelerometer data 120 and gyroscope data 121 inVERIFY_STATE 153 includes removing offset biases, filtering the data,and rotating the resulting vector data through a rotation matrix inorder to align the vector axes with the perceived direction of gravityand airplane forward acceleration. The linear acceleration and rotationrate signals are then each separately integrated to yield linearvelocity and angle data. The acceleration, velocity, rotation rate,rotation angle data are then checked to see if they are in acceptableranges of data which is characteristic of an aircraft takeoff.

FIG. 4 shows the processing steps taken by algorithm 125 in VERIFY_STATE153. Vector acceleration data A 120 and rotation rate data R 121 arereceived as well as the saved gravity vector G_takeoff 161 andhorizontal acceleration A_horiz_takeoff 162 from GROUND_STATE 151.First, rotation matrix 199 is constructed from the saved G_takeoff 161and A_horiz_takeoff 162. More specifically, rotation matrix 199 isconstructed by concatenating G_takeoff 161, with normalized horizontalacceleration A_horiz_takeoff, and the vector cross product of G_takeoff161 and normalized A_horiz_takeoff. In pseudocode the rotation matrix isexpressed as:

Rot. matrix = [G_takeoff;       A_horiz_takeoff / |A_horiz_takeoff|;      G_takeoff × (A_horiz_takeoff / |A_horiz_takeoff|) ]

By multiplying an acceleration vector by rotation matrix 199, theacceleration vector is reoriented such that the z axis is now alignedwith gravity vector G_takeoff 161, and the y-axis is aligned in thedirection of the horizontal acceleration vector A_horiz_takeoff 162. Forexample, as shown in FIG. 4, acceleration data A 120 is vectormultiplied 201 by rotation matrix 199 to produce vector Abody 202. Thez-axis vector component of Abody 202, Abody_z, represents accelerationin the direction of the perceived real world gravity. Similarly, Abody_yrepresents the acceleration or braking of the aircraft in the directionalong the runway. In summary, rotation matrix 199 is used to account forthe fact that cargo 114 may be secured in any orientation on a vehicleand the Abody vector will always be rotated such that Abody_z is in thedirection of gravity, and Abody_y is in the direction of horizontaltakeoff acceleration.

Similar to the processing to calculate Abody 202, rotation rate data R121 is rotated through vector multiplication 203 with rotation matrix199 to produce vector product 204. Vector product 204 is then passedthrough body filter 205 to produce Rbody 206. Body filter 205 is a firstorder Butterworth low pass filter with a frequency cutoff of about 0.4Hz. Rbody 206 is a low pass filtered version of R 121 in which the axeshave been rotated to correspond to the perceived aircraft body frame ofreference calculated from G_takeoff 161 and A_horiz_takeoff 162 as wasdone in calculating Abody 202. More specifically, Rbody_z represents therotation rate about the perceived aircraft yaw axis. Similarly, Rbody_yrepresents the rotation rate about the perceived aircraft pitch axis,and Rbody_x represents the rotation rate about the perceived aircraftroll axis.

Averaging block 210 maintains a running sum of acceleration data 120which is divided by the received sample count to produce vector output211. Vector output 211 is then rotated by vector multiplication 212 withrotation matrix 199 to produce Abias_body 213. Abias_body represents theraw time averaged acceleration data rotated to be aligned with theperceived aircraft frame.

Similarly, averaging block 215 maintains a running sum of rotation ratedata 121 which is divided by the count of the number of received samplesto create vector product 216. Vector product 216 is then rotated 217 byrotation matrix 199 to produce Rbias_body 213. The data used to computeAbias_body 213 and Rbias_body is limited to a fixed duration lastingabout five seconds.

Rbias_body 218 is subtracted 219 from Rbody 206 and integrated 221 toproduce angular displacement θ_body 222. Angular displacement θ_body(θ_body_x, θ_body_y, θ_body_z), 222 is vector that represents anestimate of the angle that the perceived aircraft has rotated during thetime after collection of the data for Rbias_body 218 ended. For example,θ_body_y represents the aircraft angular displacement about the pitchaxis, θ_body_z represents the aircraft angular displacement about theyaw axis, and θ_body_x represents the aircraft angular displacementabout the roll axis.

In order to determine the change in velocity, a similar integration isperformed on the Abody signal. As shown in FIG. 4, Abias_body 213 issubtracted 225 from Abody 202 to produce vector result 227. Vectorresult 227 is integrated 229 to produce velocity change vector V_body230. V_body is made up of the three components, V_body_x, V_body_y,V_body_z, each representing the change in velocity in one dimension. Forexample, V_body_y represents the change in velocity along the runwaydirection.

The rate data 121 is also passed through rate filter 233 which is afirst order Butterworth low pass filter with a cutoff frequency of about0.1 Hz. The output of rate filter 233 is R_lowpass 235.

In state VERIFY_STATE 153, a series of boundary checks are performed onthe computed acceleration, velocity change, rotation rate, and angulardisplacement variables.

The boundary checks include a check that the absolute value ofR_low_pass 235 remains<RATES_THRESHOLD 173. This is to reject times whencargo 114 is rotated too fast to be characteristic of an aircrafttakeoff. Similarly, the absolute value of Abody 211 is verified toremain greater than BODY_MAC_ACCEL. Additionally, algorithm 125 verifiesother parameters including:

abs(θ_body.y) < -MAX_PITCH_DOWN 175 θ_body.x > YAW_ROLL_LIMIT 176abs(θ_body.z) > YAW_ROLL_LIMIT 177 abs(V_body.y) > YZ_MAX_V 178abs(V_body.z) > YZ_MAX_V 178 abs(Rbody) > MAX_RATES 179 abs(Abody) >BODY_MAX_ACCEL 174

If any of the tested variables are not within acceptable range, thestate will be changed back to GROUND_STATE. However, if none of thetested variables are out of range for a duration of time greater thanVERIFY_DURATION and the pitch displacement θ_body.y is greater thanMIN_PITCH_UP and the change in runway oriented velocity V_change.x isgreater than X_MIN_V, then the state is changed to FLYING_STATE 155.

In FLYING_STATE 155 GYRO_ON/OFF_OUTPUT and RADIO_ON/OFF_OUTPUT are bothmade OFF. Also in FLYING_STATE 155 Abody is calculated using the samemethod as in VERIFY_STATE 153. A landing event is detected by sensingwhen Abody.y<−LANDING_THRESHOLD. When a landing event is detected thestate is changed back to GROUND_STATE.

Other embodiments of flight status detection system 110 include bothalgorithmic modifications and the use of additional sensor data types.

In a second embodiment, FLYING_STATE processing is augmented to involveconstant monitoring of expected aircraft maneuvers. These maneuverinclude:

a. Coordinated turns

-   -   i. Where it is expected that a horizontal plane circular motion        constraint is maintained.

b. Coupled roll and yaw rates

-   -   i. Where it is expected that roll rate leads yaw rate into a        turn, followed by constant yaw rate motion, and finally, roll        rate leads yaw rate out of a turn with opposite sign.

c. Sustained non-zero attitudes

-   -   i. Where it is expected that a climb or descent phase of flight        will maintain a pitch up or pitch down aircraft orientation for        a significant time.

A third embodiment includes a motion lockout mechanism implemented inthe GROUND_STATE to take advantage of the fact that tilts and rotationsdue to cargo handling are a strong indicator that the cargo is not aboutto be on an aircraft ready for takeoff. Thus, in the event that handlingtilts and/or rotations are detected, state transitions out of theGROUND_STATE are prevented for a duration of about two minutes.

A fourth embodiment flight status detection system includes usingpressure based confirmation of flight status as shown in FIG. 5. Aninput from pressure sensor 501 is added for in pressure based decisionmaking An absolute comparison of pressure, and/or a comparison of thepressure derivative against a threshold is used as confirmation oftakeoff. The pressure and/or pressure derivative drop below or above athreshold during aircraft events of takeoff climb and landing descent.More specifically, the pressure sensor is configured to sample thepressure at a rate of 1 Hz. This pressure sensor data is used tocalculate a time derivative of the pressure which is also low passfiltered over time. An efficient method of simultaneously low passfiltering and calculating the pressure derivative involves taking thepressure data, low pass filtering through two separate low pass filters(each with a different time constant), and then taking the differencebetween the two. This low pass filtered time derivative pressure signalis then used by the algorithm to detect aircraft climbing just aftertakeoff, and aircraft descent prior to landing.

The standard algorithm state transitions (i.e. FIG. 2) are then verifiedby detecting changes in the smoothed time derivative pressure signal.More concretely, a transition from VERIFY_TAKEOFF to FLYING_STATE stateas shown in FIG. 2, is verified when the smoothed time derivativepressure signal drops to less than a TAKEOFF_PRESSURE_THRESHOLD.Similarly, a transition from FLYING_STATE to GROUND_STATE is verifiedwhen the smoothed time derivative pressure signal becomes greater than aDESCENDING_PRESSURE_THRESHOLD. The embodiment can utilize a statemachine specifically for pressure signal processing (FIG. 5) which isseparate from the standard algorithm state machine (FIG. 2). In thisconfiguration, the pressure signal processing state machine 510 runs inparallel to the a state machine as described in the first embodiment andshown in FIG. 2.

As shown in FIG. 5, pressure state machine 510 having the states ofWAITING_STATE 520, TAKEOFF_STATE 525, CLIMB_FINISHED_STATE 527,DESCENDING_STATE 529, and LANDED_STATE 520. The state machine begins inWAITING_STATE 520. If the smoothed time derivative pressure signal everhas a value below the negative value of threshold PTHRESH, the statewill transition to TAKE_OFF_STATE 525 indicating that there is alikelihood that the aircraft is climbing. Similarly, while inWAITING_STATE 520, if the smoothed time derivative pressure signal everhas a value above threshold PTHRESH, the state will transition toDESCENDING_STATE 529 indicating that there is a likelihood that theaircraft is descending.

While in TAKEOFF_STATE 525, if the absolute value of the smoothed timederivative pressure signal ever drops below threshold PTHRESH, the statewill change to CLIMB_FINISHED_STATE 527. In CLIMB_FINISHED_STATE 527 amessage indicating the climb has finished may be communicated to otherportions of the algorithm, then the state will go back to WAITING_STATE520. Similarly, while in DESCENDING_STATE 525, if the absolute value ofthe smoothed time derivative pressure signal ever drops below thresholdPTHRESH, the state will change to LANDED_STATE 531. In LANDED_STATE 531a message indicating the aircraft has landed may be communicated toother portions of the algorithm, then the state will go back toWAITING_STATE 520. The addition of pressure sensor 501 and the analysisof the pressure signal as described allows a secondary means ofdetermining flight status, which may be used as a backup method or as away of confirming flight state as determined by other methods.

In a fifth embodiment, a failsafe timer is implemented on the flightstatus detection system. When in flight, the timer is used to detectwhen a total possible flight time threshold is passed to bring the statemachine out FLYING_STATE to ensure failsafe operation for missed landingdetections.

In a sixth embodiment, geographic location data is used to implement ageographic fencing. A library of airport locations can be comparedagainst real-time location of the device determined by GPS, celltriangulation, or WiFi. When the system GPS indicates the apparatus iswithin an airport boundary, the takeoff detection algorithm can be runwith more aggressive parameters. For example, higher accelerometersampling rates may be used, and/or the gyroscope may be turned oninstead of waiting for a takeoff acceleration. When the system is notwithin an airport boundary, it could run in a less aggressive mode, ornot at all to save device power and extend battery life.

In a seventh embodiment, the audio signature of a jet or propellerengine could also be detected with a microphone and DSP signalprocessing to further confirm the device is within an aircraft andshould turn its radios off.

In an eighth embodiment, a magnetometer is used to help detect magneticfields and field signatures in proximity to the cargo. For example, theunique magnetic signature of various aircraft bodies could be detectedwith a magnetometer and DSP signal processing to further confirm thedevice is within an aircraft and should turn its radios off.

The disclosed flight status detection system and methods resulted inseveral advantages and surprising results. The system and method wasable to accurately distinguish between aircraft flight events and othertransportation movements resulting in a system and method capable ofsafely shutting down an external radio before flight takeoff and turningthe radio back on soon after landing. False radio shutoffs are alsominimized with the disclosed system and method. Further, the method andsystem is capable of operating accurately while using only low samplerates and processing power. This results in significant power savingsand allows the system and method to work for extended periods withoutrecharging or new batteries. Additionally, while the disclosed deviceand method can have a GPS, a GPS is not necessary for accuratefunctioning. Usage of the method and device without a GPS offerssubstantial power savings over a device or method which uses a GPS. Mostimportantly, perhaps, the disclosed method and system offer the benefita completely self-contained system that does not require external inputsor coordination with other systems. This allows deployment withoutdependencies on external infrastructure and offers high reliability thatis independent of other system failures.

While several embodiments of the flight status detector system has beenshown and described, and several modifications thereof discussed,persons skilled in this art will readily appreciate that variousadditional changes may be made without departing from the spirit of theinvention.

1. A low power method for determining whether a cargo destined for airtransport is in a flying state comprising the steps of: providing ahousing for attachment to a cargo, said housing comprising: anaccelerometer for detecting a linear acceleration, a gyroscope fordetecting an angular rate and a controller; measuring a linearacceleration with said accelerometer; measuring an angular rate withsaid gyroscope; providing said measured linear acceleration and angularrate to said controller; and generating a flight status output signalindicating whether said housing is in a flying state as a function ofsaid linear acceleration signal and said angular rate signal.
 2. Themethod as set forth in claim 1, and further comprising the step ofshutting off a radio as a function of said flight status output signal.3. The method as set forth in claim 1, wherein said measured linearacceleration comprises a three independent axis linear accelerationsignal and said measured angular rate comprises a three independent axisrotation rate signal.
 4. (canceled)
 5. (canceled)
 6. The method as setforth in claim 1, wherein said flight status output signal comprises anin-flight status state and an on-the-ground status state.
 7. (canceled)8. The method as set forth in claim 6, and further comprising the stepof providing a radio turn off signal when said flight status outputsignal comprises an in-flight status state.
 9. The method as set forthin claim 1, wherein said function is a state machine algorithmcomprising the steps of filtering said measured linear acceleration toproduce a filtered linear acceleration, filtering said measured angularrate to produce a filtered angular rate, and changing a state as afunction of said filtered linear acceleration and filtered angular rate.10. The method as set forth in claim 1, and further comprising the stepof powering said gyroscope as a function of said flight status outputsignal such that said gyroscope is powered off when said flight statusoutput signal comprises an on-the-ground status state and said gyroscopeis powered on when said flight status output signal comprises averify-takeoff status state.
 11. (canceled)
 12. (canceled) 13.(canceled)
 14. The method as set forth in claim 1, wherein said step ofgenerating a flight status output signal comprises the steps of removinga gravitational component of said measured linear acceleration with ahigh pass filter having a cutoff frequency of about 0.01 Hz and savingsaid removed gravitational component.
 15. (canceled)
 16. (canceled) 17.(canceled)
 18. The method as set forth in claim 1, wherein said step ofgenerating a flight status output signal further comprises the step ofdetecting a linear acceleration characteristic of an aircraft takeoff.19. The method as set forth in claim 18, wherein said linearacceleration characteristic of a takeoff is an acceleration signal witha frequency between 0.01 Hz and 0.1 Hz and a magnitude of between about0.15 g and 0.5 g maintained for a duration of about two seconds. 20.(canceled)
 21. (canceled)
 22. The method as set forth in claim 18, andfurther comprising the step of saving a takeoff acceleration of saidlinear acceleration signal when said linear acceleration characteristicof an aircraft takeoff is detected and wherein said saved takeoffacceleration comprises a windowed average.
 23. (canceled)
 24. The methodas set forth in claim 18, wherein said step of generating a flightstatus output signal further comprises the step of changing from anon-the-ground state to a verify-takeoff state when said linearacceleration characteristic of an aircraft takeoff is detected.
 25. Themethod as set forth in claim 22, and further comprising the step ofsaving a gravitational acceleration component from said linearacceleration signal and wherein said gravitational acceleration is savedfrom a time period before a transition to a verity-takeoff state. 26.(canceled)
 27. (canceled)
 28. The method as set forth in claim 18,wherein said step of generating a flight status output signal furthercomprises the step of detecting an angular rate characteristic of anaircraft takeoff.
 29. The method as set forth in claim 28, wherein saidangular rate characteristic of an aircraft takeoff is an angular rate inwhich an x-axis angular rate, a y-axis angular rate, and a z-axisangular rate are all less than an angular rate limit threshold.
 30. Themethod as set forth in claim 28, wherein said angular rate istransformed through a rotation, said rotation configured and arranged toreorient said measured angular rate into a yaw rate, a roll rate, and apitch rate.
 31. The method as set forth in claim 30, wherein saidrotation is a cross product of said saved gravitational acceleration andsaid saved takeoff acceleration.
 32. The method as set forth in claim30, and further comprises the steps of integrating said angular ratewith respect to time to produce a pitch angle displacement, a yaw angledisplacement, and a roll angle displacement and setting said flightstatus output signal to an on-the-ground state when said pitch angledisplacement, yaw angle displacement, or roll angle displacement aregreater than an angle displacement threshold.
 33. (canceled)
 34. Themethod as set forth in claim 33, and further comprising the steps of:providing a pressure sensor; detecting when said pressure sensor has apressure decrease rate less than a takeoff pressure decrease threshold;and detecting when said pressure sensor has a pressure increase rategreater than a landing pressure increase threshold.
 35. (canceled) 36.(canceled)
 37. A system for detecting an aircraft flight statuscomprising: an accelerometer for detecting a linear acceleration andhaving a linear-acceleration output signal; a gyroscope for detecting anangular rate and having an angular rate output signal; and a controllerconfigured and arranged to produce a flight status output signal as afunction of said linear acceleration output signal and said angular rateoutput signal, said flight status output signal comprising an in-flightstate and second state; and wherein said accelerometer, gyroscope andcontroller are in a housing.
 38. The system as set forth in claim 37,and wherein said controller is further configured to shut off a radio asa function of said flight status output signal.
 39. The system as setforth in claim 37, wherein said linear acceleration output signalcomprises a three independent axis linear acceleration signal and saidangular rate output signal comprises a three independent axis rotationrate signal.
 40. (canceled)
 41. (canceled)
 42. The system as set forthin claim 37, wherein said second state of said flight status outputsignal comprises an on-the-ground status state.
 43. The system set forthin claim 42, wherein said in-flight status state comprises averify-takeoff status state.
 44. (canceled)
 45. (canceled) 46.(canceled)
 47. The method as set forth in claim 43, wherein saidcontroller is configured and arranged to power off said gyroscope whensaid flight status output signal comprises said an on-the-ground statusstate and to power on said gyroscope when said flight status outputsignal comprises said verify-takeoff status state.
 48. (canceled) 49.The system as set forth in claim 37, wherein said function comprises afilter configured and arranged to produce a filtered linearacceleration, and a filter configured and arranged to produce a filteredangular rate.
 50. The system as set forth in claim 37, wherein saidcontroller is configured and arranged to remove a gravitationalcomponent of said linear acceleration output signal, said system furthercomprises a high pass filter configured and arranged to remove saidgravitational component from said linear acceleration output signal,wherein said high pass filter has a cutoff frequency of about 0.01 Hz,and wherein said controller is configured and arranged to save saidremoved gravitational component.
 51. (canceled)
 52. (canceled) 53.(canceled)
 54. The system as set forth in claim 37, wherein saidcontroller is configured and arranged to detect a linear accelerationcharacteristic of an aircraft takeoff.
 55. The system as set forth inclaim 54, wherein said linear acceleration characteristic of takeoff isan acceleration signal with a frequency between 0.01 Hz and 0.1 Hz and amagnitude of between about 0.15 g and 0.5 g maintained for a duration ofabout two seconds.
 56. (canceled)
 57. (canceled)
 58. The system as setforth in claim 54, wherein said controller is configured and arranged tosave said acceleration output signal when a linear accelerationcharacteristic of an aircraft is detected.
 59. (canceled)
 60. The systemas set forth in claim 58, wherein said controller is configured andarranged to change said flight status output signal from anon-the-ground state to a verify-takeoff state when said linearacceleration characteristic of an aircraft takeoff is detected.
 61. Thesystem as set forth in claim 58, wherein said controller is configuredand arranged to save a gravitational acceleration component from saidlinear acceleration signal and wherein said gravitational accelerationis saved from a time period before a transition to a verify-takeoffstate.
 62. (canceled)
 63. (canceled)
 64. The system as set forth inclaim 58, wherein said controller is configured and arranged to detectan angular rate characteristic of an aircraft takeoff.
 65. The system asset forth in claim 64, wherein said angular rate characteristic of anaircraft takeoff is an angular rate in which an x-axis angular rate, ay-axis angular rate, and a z-axis angular rate are all less than anangular rate limit threshold.
 66. The system as set forth in claim 64,and further comprising a rotation transformer configured and arranged totransformed said angular rate output signal through a rotation, saidrotation configured and arranged to reorient said measured angular rateinto a yaw rate, a roll rate, and a pitch rate.
 67. The system as setforth in claim 66, wherein said rotation is a cross product of saidsaved gravitational acceleration and said saved takeoff acceleration.68. The system as set forth in claim 66, and further comprising anintegrator configured and arranged to integrate said angular rate withrespect to time to produce a pitch angle displacement, a yaw angledisplacement, and a roll angle displacement and wherein said controlleris configured and arranged to set said flight status output signal to anon-the-ground state when said pitch angle displacement, yaw angledisplacement, or roll angle displacement are greater than an angledisplacement threshold.
 69. (canceled)
 70. The system as set forth inclaim 69, and further comprising a pressure sensor.
 71. The system asset forth in claim 70, and wherein said controller is configured andarranged to detect when said pressure sensor has a pressure decreaserate less than a takeoff pressure decrease threshold and to detect whensaid pressure sensor has a pressure increase rate greater than a landingpressure increase threshold.
 72. (canceled)
 73. (canceled) 74.(canceled)